Gas turbine engine geared architecture axial retention arrangement

ABSTRACT

A gas turbine engine includes a fan, an engine static structure, a geared architecture to drive the fan and supported relative to the static structure, a fan drive turbine to drive the geared architecture, a first member secured to the geared architecture, and a second member secured to the engine static structure and configured to cooperate with the first member to limit movement of the geared architecture relative to the static structure. A fan drive gear system and method are also disclosed.

This application is a continuation application of U.S. patentapplication Ser. No. 13/435,353, which was filed on Mar. 30, 2012.

BACKGROUND

This disclosure relates to limiting axial movement of a gearedarchitecture within a turbomachine during an extreme event.

Turbomachines, such as gas turbine engines, typically include a fansection, a turbine section, a compressor section, and a combustorsection. Turbomachines may employ a geared architecture connecting thefan section and the turbine section.

Support structures are used to hold the geared architecture within theturbomachine. The support structures may be relatively compliant toaccommodate some movement of the geared architecture relative to otherportions of the turbomachine. Extreme engine events such as fan bladeloss or failure of fan shaft bearing supports may encourage significantaxial movement of the geared architecture and the fan, relative to otherportions of the turbomachine. These movements are undesirable as isknown. The relatively compliant support structures may not providedesired axial retention of the geared architecture during extreme engineevents.

SUMMARY

A gas turbine engine according to an exemplary embodiment of thisdisclosure, among other possible things includes a fan, an engine staticstructure, a geared architecture to drive the fan and supported relativeto the static structure, a fan drive turbine to drive the gearedarchitecture, a first member secured to the geared architecture, and asecond member secured to the engine static structure and configured tocooperate with the first member to limit movement of the gearedarchitecture relative to the static structure.

In a further embodiment of any of the foregoing gas turbine engines, thefirst and second members are circumferentially aligned with one anotherand spaced apart from one another during a normal operating condition.

In a further embodiment of any of the foregoing gas turbine engines, thefirst and second members limit axial movement of the geared architecturerelative to the engine static structure.

In a further embodiment of any of the foregoing gas turbine engines,includes a flex support supporting the geared architecture relative tothe engine static structure.

In a further embodiment of any of the foregoing gas turbine engines,includes a support structure secured to the geared architecture and theflex support. The support structure includes at least one of a torqueframe, a carrier, and a lubrication manifold, and the second member isremovably secured to at least one of the torque frame, the carrier, andthe lubrication manifold.

In a further embodiment of any of the foregoing gas turbine engines, thegas turbine engine has a bypass ratio greater than about six (6).

In a further embodiment of any of the foregoing gas turbine engines, thegeared architecture has a gear reduction ratio greater than about 2.3.

In a further embodiment of any of the foregoing gas turbine engines, thegeared architecture is a planetary gear train.

In a further embodiment of any of the foregoing gas turbine engines, theplanetary gear train includes a plurality of gears supported within thecarrier. The carrier is fixed against rotation by the torque frame. Acentral sun gear is operatively connected to the fan drive turbine. Aring gear is configured to drive the fan.

In a further embodiment of any of the foregoing gas turbine engines, thefan is operatively coupled to the geared architecture via a fan shaft,and the fan shaft is supported relative to the engine static structureby at least two bearings.

In a further embodiment of any of the foregoing gas turbine engines, thefan drive turbine has a pressure ratio greater than about five (5).

In a further embodiment of any of the foregoing gas turbine engines, theflex support includes a bellow, an annular mounting flange opposite thebellow and the first member is removably secured to the annular mountingflange.

In a further embodiment of any of the foregoing gas turbine engines, thefirst member and the second member are U-shaped brackets oriented inopposite radial positions.

In a further embodiment of any of the foregoing gas turbine engines,further includes a brace to strengthen the axial retention of the firstmember.

In a further embodiment of any of the foregoing gas turbine engines, atleast one of the torque frame and the flex support includes at least onefeature configured to limit annular rotation of at least one of thefirst and second members.

In a further embodiment of any of the foregoing gas turbine engines, thefirst and second members engage one another by axial movement inopposite directions.

A method of assembling a gas turbine engine in which a fan is driven bya speed reduction device according to an exemplary embodiment of thisdisclosure, among other possible things includes providing attachmentfeatures in a first structure and a second structure, securing first andsecond members respectively to the first and second structures, andinstalling the first structure onto an engine static structure and thespeed reduction device onto the second structure such that the first andsecond members are engageable with one another during an extreme event.

In a further embodiment of any of the foregoing methods, the firststructure is a flex support having a bellow and an annular mountingflange opposite the bellow, and the securing step includes mounting alubrication manifold onto the second structure, and securing the secondmembers over the lubrication manifold.

In a further embodiment of any of the foregoing methods, furtherincludes the step of positioning the first and second members in a firstangular position relative to one another, and rotating the first andsecond members from the first angular position to a second angularposition against a brace.

In a further embodiment of any of the foregoing methods, the first andsecond members are arranged in an axially spaced relation to one anotherin an installed condition, and are configured to engage one another bymoving axially in opposite directions.

A fan drive gear system for a gas turbine engine according to anexemplary embodiment of this disclosure, among other possible thingsincludes a geared architecture configured to drive a fan, a first membersecured to the geared architecture, and a second member configured forsecuring to a static structure and configured to cooperate with thefirst member to limit movement of the geared architecture relative tothe static structure.

In a further embodiment of any of the foregoing fan drive gear systems,the first and second members are circumferentially aligned with oneanother and spaced apart from one another during a normal operatingcondition.

In a further embodiment of any of the foregoing fan drive gear systems,includes a flex support supporting the geared architecture relative tothe static structure.

In a further embodiment of any of the foregoing fan drive gear systems,includes a support structure secured to the geared architecture and theflex support. The support structure includes at least one of a torqueframe, a carrier, and a lubrication manifold, and the second member isremovably secured to at least one of the torque frame, the carrier, andthe lubrication manifold.

In a further embodiment of any of the foregoing fan drive gear systems,the geared architecture includes a plurality of gears supported withinthe carrier. The carrier is fixed against rotation by the torque frame.A central sun gear is operatively connected to the fan drive turbine,and a ring gear configured to drive the fan.

In a further embodiment of any of the foregoing fan drive gear systems,the geared architecture has a speed reduction ratio greater than about2.3.

A method of designing a gas turbine engine in which a fan is driven by aspeed reduction device according to an exemplary embodiment of thisdisclosure, among other possible things includes defining attachmentfeatures in a first structure and a second structure, configuring firstand second members for securement respectively to the first and secondstructures, and defining the first structure for attachment to an enginestatic structure and the second structure for attachment to the speedreduction device such that the first and second members are engageablewith one another during an extreme event.

In a further embodiment of any of the foregoing methods, the firststructure is defined as a flex support having a bellow and an annularmounting flange opposite the bellow, the second structure is configuredfor securement to a lubrication manifold and the second member isconfigured for securement over the lubrication manifold.

In a further embodiment of any of the foregoing methods, furtherincludes configuring the first and second members to be positioned in afirst angular position relative to one another and that rotating thefirst and second members from the first angular position to a secondangular position abuts against a brace.

In a further embodiment of any of the foregoing methods, includesdefining the first and second members to be arranged in an axiallyspaced relation to one another in an installed condition and to engageone another by moving axially in opposite directions during the extremeevent.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 shows a partial section view of an example gas turbine engine.

FIG. 2A depicts a highly schematic view of an example gearedarchitecture support assembly of the FIG. 1 gas turbine engine duringnormal operation.

FIG. 2B depicts a highly schematic view of the FIG. 2A gearedarchitecture support during an extreme event.

FIG. 3 is a cross-sectional view of a geared architecture and an examplesupport assembly during normal operation.

FIG. 4 is a rear view of an example oil manifold of the supportassembly.

FIG. 5 is a front view of an example flex support of the supportassembly.

FIG. 6A is a front view of the oil manifold and flex support in a firstangular position during assembly.

FIG. 6B is a top elevational view of the support assembly in the firstangular position.

FIG. 7 is a front view of the oil manifold and flex support in a secondangular position after assembly.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flowpath B whilethe compressor section 24 drives air along a core flowpath C forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines including three-spool architectures.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure (or first) compressor section 44and a low pressure (or first) turbine section 46. The inner shaft 40 isconnected to the fan 42 through a geared architecture 48 to drive thefan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a high pressure(or second) compressor section 52 and high pressure (or second) turbinesection 54. A combustor 56 is arranged between the high pressurecompressor 52 and the high pressure turbine 54. A mid-turbine frame 57of the engine static structure 36 is arranged generally between the highpressure turbine 54 and the low pressure turbine 46. The mid-turbineframe 57 supports one or more bearing systems 38 in the turbine section28. The inner shaft 40 and the outer shaft 50 are concentric and rotatevia bearing systems 38 about the engine central longitudinal axis A,which is collinear with their longitudinal axes. As used herein, a “highpressure” compressor or turbine experiences a higher pressure than acorresponding “low pressure” compressor or turbine.

The core airflow C is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than ten (10), the gearedarchitecture 48 is an epicyclic gear train, such as a star gear systemor other gear system, with a gear reduction ratio of greater than about2.3 and the low pressure turbine 46 has a pressure ratio that is greaterthan about 5. In one disclosed embodiment, the engine 20 bypass ratio isgreater than about ten (10:1), the fan diameter is significantly largerthan that of the low pressure compressor 44, and the low pressureturbine 46 has a pressure ratio that is greater than about 5:1. Lowpressure turbine 46 pressure ratio is pressure measured prior to inletof low pressure turbine 46 as related to the pressure at the outlet ofthe low pressure turbine 46 prior to an exhaust nozzle. It should beunderstood, however, that the above parameters are only exemplary of oneembodiment of a geared architecture engine and that the presentinvention is applicable to other gas turbine engines including directdrive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned per hour divided by lbf of thrustthe engine produces at that minimum point. “Fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tambient degR)/518.7)̂0.5]. The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

A fan shaft 60 interconnects the geared architecture 48 to the fan 42.The fan shaft 60 is supported by a pair of bearings 38, which aretapered roller bearings in one example. The bearings normally limit theaxial travel of the fan shaft 60 and fan 42. During operation, theengine 20 may experience extreme events such as a fan blade loss or afailure of a fan shaft bearing support 62 supporting the bearings 38,which is part of the engine static structure 36. In such events, the fan42 may undesirably tend to move axially forward relative to the otherportions of the engine 20, such that the fan 42 and associatedcomponents could become disengaged from the engine 20.

Referring to FIGS. 2A and 2B with continuing reference to FIG. 1, theexample engine 20 includes other features that limit movement of the fan42 during an extreme event, particularly if the bearings 38 or bearingsupport 62 are ineffective. For example, the example engine 20 includesa geared architecture support assembly 64 that limits forward movementof the fan 42 and the geared architecture 48 during an extreme event.

The example support assembly 64 includes at least a first member 66 anda second member 68. The first member 66 and the second member 68 arerespectively operatively connected to the geared architecture 48 and theengine static structure 36, which functions as a mechanical ground. Acompliant flex support 70 mounts the geared architecture 48 to theengine static structure 36. In the example, the first member 66 and thesecond member 68 are both arranged axially aft the geared architecture48 relative to a direction of flow through the engine 20.

During normal engine operation, the first and second members 66, 68 arespaced apart from one another providing a gap 72, as shown in FIG. 2A.During an extreme event, such as a blade loss, the geared architecture48 may experience an extreme load in the direction F due to the fan 42rotating and attempting to move axially forward relative to otherportions of the engine 20, as shown in FIG. 2B. In such an event,extreme movements of the geared architecture 48 are limited by thecooperation of the first and second members 66, 68 such that the loadingin the direction F causes the first and second members 66, 68 to engageone another at area 74. This contact blocks movement of the gearedarchitecture 48 axially forward. Since the geared architecture 48 isconnected to the fan 42, limiting movement of the geared architecture 48may prevent the fan 42 from moving axially forward the remainingattached portions of the engine 20.

One example support assembly 64 is illustrated schematically in FIG. 3.The flex support 70 is secured to a carrier 81 by a torque frame 82. Alubrication manifold 78 is arranged axially between the carrier 81 andthe flex support 70. A geartrain 84 of the geared architecture 48 issupported by the torque frame 82. In one example, the geartrain 84 is aplanetary gear arrangement in which planetary gears are supported by thecarrier 81 and fixed against rotation by torque frame 82. A central sungear receives rotational drive from the inner shaft 40 (FIG. 1) and aring gear rotationally drives the fan 42 through the fan shaft 60 (FIG.1).

The flex support 70 includes a bellow 89, which is provided by a wallthat doubles back on itself to provide an undulation. The bellow 89accommodates a relative movement of the geared architecture 48 relativeto the engine static structure 36. An annular mounting flange 91 at anend opposite the bellow 89 is rotationally fixed relative to the enginestatic structure 36 by fasteners, splines and/or other means.

In the example illustrated, the support assembly 64 is provided by setsof first and second members 66, 68, which are removably securedrespectively to the flex support 70 and the lubrication manifold 78. Thefirst and second members 66, 68 are provided by U-shaped bracketsoriented in opposite radial positions from one another to facilitateassembly. In the example, each set of members include fivecircumferentially spaced brackets.

The support assembly 64 may be retrofitted to existing gas turbineengines with geared architectures. In one example, first and secondmachined surfaces 86, 88 are respectively provided on the flex support70 and a back side 87 of the lubrication manifold 78. If these machinedsurfaces are not provided on existing parts, the manufacturer can millthese surfaces, for example, as part of the retrofitting process. Firstfasteners 90 secure the first member 66 to the end 91. Second fasteners92 secure the second member 68 to the lubrication manifold 78 andcarrier 81. Existing geared architectures may be retrofitted byreplacing the pre-existing fasteners that secure the lubricationmanifold 78 to the carrier 81 with longer fasteners while reusingexisting holes in the carrier 81 and the lubrication manifold 78. Thefirst and second fasteners 90, 92 are threaded fasteners in one example.

Each first member 66 is provided by spaced apart legs 94 joined by abend 96. Similarly, each second member 68 is provided by spaced apartlegs 98 joined by a bend 100. The legs 94, 98 are axially spaced fromone another to provide the gap 72 during normal operation.

Referring to FIG. 4, the lubrication manifold 78 may include integrallyformed lubrication passages 79 that are cast into the lubricationmanifold 78 to provide a unitary structure. The second members 68include a second flange 104 supporting a leg 98 and secured to an outerperiphery of the lubrication manifold. Alternately, the second member 68could be secured directly to the carrier 81 through windows or scallops76 in the lubrication manifold 78. Referring to FIG. 5, the first member66 includes a first flange 102 supporting a leg 94 and secured to theannular mounting flange 91 of the flex support 70.

Referring to FIGS. 6A-7, with the first and second members 66, 68respectively are secured to the flex support 70 and the lubricationmanifold 78. The lubrication manifold 78 is arranged in a first angularposition 106, illustrated in FIGS. 6A and 6B, such that the first andsecond members 66, 68 are circumferentially adjacent to one another butmisaligned. The lubrication manifold 78 is rotated relative to the flexsupport 70 to circumferentially align the first and second members 66,68 relative to one another in a second angular position 108. In thisexample, the first member 66 includes a brace 110 to strengthen theaxial retention of the member. The annular rotation is limited byexisting features, which are machined into the torque frame 82 and theflex support 70.

Features of the disclosed examples include a support structure thatpermits some movement of a geared architecture relative to otherportions of an engine during normal operation of the engine, but limitsmovements during extreme events, particularly axially forward movementsof the geared architecture.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. A gas turbine engine comprising: a fan; an enginestatic structure; a geared architecture to drive the fan and supportedrelative to the static structure; a fan drive turbine to drive thegeared architecture; a first member secured to the geared architecture;and a second member secured to the engine static structure andconfigured to cooperate with the first member to limit movement of thegeared architecture relative to the static structure.
 2. The gas turbineengine of claim 1, wherein the first and second members arecircumferentially aligned with one another and spaced apart from oneanother during a normal operating condition.
 3. The gas turbine engineof claim 2, wherein the first and second members limit axial movement ofthe geared architecture relative to the engine static structure.
 4. Thegas turbine engine of claim 3, including a flex support supporting thegeared architecture relative to the engine static structure.
 5. The gasturbine engine of claim 4 including a support structure secured to thegeared architecture and the flex support, wherein the support structurecomprises at least one of a torque frame, a carrier, and a lubricationmanifold, and the second member is removably secured to at least one ofthe torque frame, the carrier, and the lubrication manifold.
 6. The gasturbine engine of claim 4, wherein the gas turbine engine has a bypassratio greater than about six (6).
 7. The gas turbine engine of claim 6,wherein the geared architecture has a gear reduction ratio greater thanabout 2.3.
 8. The gas turbine engine of claim 7, wherein the gearedarchitecture is a planetary gear train.
 9. The gas turbine engine ofclaim 8, wherein the planetary gear train includes: a plurality of gearssupported within the carrier, the carrier is fixed against rotation bythe torque frame; a central sun gear is operatively connected to the fandrive turbine; and a ring gear configured to drive the fan.
 10. The gasturbine engine of claim 9, wherein the fan is operatively coupled to thegeared architecture via a fan shaft, and the fan shaft is supportedrelative to the engine static structure by at least two bearings. 11.The gas turbine engine of claim 10, wherein the fan drive turbine has apressure ratio greater than about five (5).
 12. The gas turbine engineof claim 11, wherein the flex support includes a bellow, an annularmounting flange opposite the bellow and the first member is removablysecured to the annular mounting flange.
 13. The gas turbine engine ofclaim 12, wherein the first member and the second member are U-shapedbrackets oriented in opposite radial positions.
 14. The gas turbineengine of claim 13, further including a brace to strengthen the axialretention of the first member.
 15. The gas turbine engine of claim 14,wherein at least one of the torque frame and the flex support includesat least one feature configured to limit annular rotation of at leastone of the first and second members.
 16. The gas turbine engine of claim15, wherein the first and second members engage one another by axialmovement in opposite directions.
 17. A method of assembling a gasturbine engine in which a fan is driven by a speed reduction device, themethod comprising the steps of: providing attachment features in a firststructure and a second structure; securing first and second membersrespectively to the first and second structures; and installing thefirst structure onto an engine static structure and the speed reductiondevice onto the second structure such that the first and second membersare engageable with one another during an extreme event.
 18. The methodaccording to claim 17, wherein the first structure is a flex supporthaving a bellow and an annular mounting flange opposite the bellow, andthe securing step includes mounting a lubrication manifold onto thesecond structure, and securing the second members over the lubricationmanifold.
 19. The method according to claim 18, further including thestep of positioning the first and second members in a first angularposition relative to one another, and rotating the first and secondmembers from the first angular position to a second angular positionagainst a brace.
 20. The method according to claim 19, wherein the firstand second members are arranged in an axially spaced relation to oneanother in an installed condition, and are configured to engage oneanother by moving axially in opposite directions.
 21. A fan drive gearsystem for a gas turbine engine comprising: a geared architectureconfigured to drive a fan; a first member secured to the gearedarchitecture; and a second member configured for securing to a staticstructure and configured to cooperate with the first member to limitmovement of the geared architecture relative to the static structure.22. The fan drive gear system as recited in claim 21, wherein the firstand second members are circumferentially aligned with one another andspaced apart from one another during a normal operating condition. 23.The fan drive gear system as recited in claim 22, including a flexsupport supporting the geared architecture relative to the staticstructure.
 24. The fan drive gear system as recited in claim 23,including a support structure secured to the geared architecture and theflex support, wherein the support structure comprises at least one of atorque frame, a carrier, and a lubrication manifold, and the secondmember is removably secured to at least one of the torque frame, thecarrier, and the lubrication manifold.
 25. The fan drive gear system asrecited in claim 24, wherein the geared architecture includes: aplurality of gears supported within the carrier, the carrier is fixedagainst rotation by the torque frame; a central sun gear is operativelyconnected to the fan drive turbine; and a ring gear configured to drivethe fan.
 26. The fan drive gear system as recited in claim 21, whereinthe geared architecture has a speed reduction ratio greater than about2.3.
 27. A method of designing a gas turbine engine in which a fan isdriven by a speed reduction device, the method, comprising the steps of:defining attachment features in a first structure and a secondstructure; configuring first and second members for securementrespectively to the first and second structures; and defining the firststructure for attachment to an engine static structure and the secondstructure for attachment to the speed reduction device such that thefirst and second members are engageable with one another during anextreme event.
 28. The method according to claim 27, wherein the firststructure is defined as a flex support having a bellow and an annularmounting flange opposite the bellow, the second structure is configuredfor securement to a lubrication manifold and the second member isconfigured for securement over the lubrication manifold.
 29. The methodaccording to claim 27, further including configuring the first andsecond members to be positioned in a first angular position relative toone another and that rotating the first and second members from thefirst angular position to a second angular position abuts against abrace.
 30. The method according to claim 27, including defining thefirst and second members to be arranged in an axially spaced relation toone another in an installed condition and to engage one another bymoving axially in opposite directions during the extreme event.